I find it difficult to accept the idea that at 3000-4000 Kelvin superheatedCO2, NO2 and or H20 do not behave like a plasma. I would think that highly agitated electrons and nuclei would not tend to recombine immediately inside rocket motor combustion chambers; but I could be wrong.

The chief difference between chemical engines and plasma engines us that chemical engines do not interact with the exhaust gases. All its thrust comes from the expansion of gases. Plasma engines use electric and magnetic fields to accelerate the plasma exhaust gases.

Chris Whoever loves correction loves knowledge, but he who hates reproof is stupid.

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The chief difference between chemical engines and plasma engines us that chemical engines do not interact with the exhaust gases. All its thrust comes from the expansion of gases. Plasma engines use electric and magnetic fields to accelerate the plasma exhaust gases.

I think we're not on the same page; but that's my fault.My question was sort of ambivalent...my regrets.But I got the answer I was looking for below.

Weak plasma or strong. A combustion generated plasma rather than an electrically generated oneis what I would call it.Thanks for your answer. Cheers.BTW...a 'weak' combustion generated plasma still presents unique possibilities if one thinks outside the box....But I'll take that topic to"advanced concepts".

And JP Aerospace suggests a plasma engine in which the plasma is the exhaust of a chemical engine, which is then accelerated by electromagnetic means. The advantage of this weird hybrid approach is that it's not necessary to generate a large amount of electricity to create the plasma in the first place.

FYI, a low temperature plasma is normally used to heat treat components in Plasma nitriding; the combination of low pressure and DC high voltage generates a Hydrogen- Nitrogen plasma acceleration layer (the glowing layer in pic).Concerning rockets engines, beb's explanation is the best, plasma sensibility to EM fields is used to accelerate it (MHD generators do the inverse job, picking energy from the plasma content of a very high temperature combustion flame).

Where do you see it as a constraint? As free molecular heating?For Pegasus, it is not just the fairing. The wings char.Also, there is interference heating.

Well I just had one study that claimed that heat flux is an important constraint when simulating launcher trajectories. However, I cannot find a second source, so perhaps my assumption is wrong?

A second source would be the Atlas V User Guide:"For Atlas V 500 series missions, the PLF is jettisoned during the booster phase of flight. Before PLF jettison,the RD-180 engine is throttled down to maintain 2.5 g acceleration. Typically, the PLF is jettisoned when the3-sigma free molecular heat flux falls below 1,135 W/m2 (360 Btu/ft2-hr). For sensitive SC, PLF jettison canbe delayed to reduce the heat flux with minor performance loss. After PLF jettison, the RD-180 is throttledup."So it can be a constraint that really originates from the payload and not necessarily from the launch vehicle.

Where do you see it as a constraint? As free molecular heating?For Pegasus, it is not just the fairing. The wings char.Also, there is interference heating.

Well I just had one study that claimed that heat flux is an important constraint when simulating launcher trajectories. However, I cannot find a second source, so perhaps my assumption is wrong?

A second source would be the Atlas V User Guide:"For Atlas V 500 series missions, the PLF is jettisoned during the booster phase of flight. Before PLF jettison,the RD-180 engine is throttled down to maintain 2.5 g acceleration. Typically, the PLF is jettisoned when the3-sigma free molecular heat flux falls below 1,135 W/m2 (360 Btu/ft2-hr). For sensitive SC, PLF jettison canbe delayed to reduce the heat flux with minor performance loss. After PLF jettison, the RD-180 is throttledup."So it can be a constraint that really originates from the payload and not necessarily from the launch vehicle.

The assumption is wrong and therefore that isn't the second source.It only affects fairing jettison time. The trajectory isn't really affected.

Where do you see it as a constraint? As free molecular heating?For Pegasus, it is not just the fairing. The wings char.Also, there is interference heating.

Well I just had one study that claimed that heat flux is an important constraint when simulating launcher trajectories. However, I cannot find a second source, so perhaps my assumption is wrong?

A second source would be the Atlas V User Guide:"For Atlas V 500 series missions, the PLF is jettisoned during the booster phase of flight. Before PLF jettison,the RD-180 engine is throttled down to maintain 2.5 g acceleration. Typically, the PLF is jettisoned when the3-sigma free molecular heat flux falls below 1,135 W/m2 (360 Btu/ft2-hr). For sensitive SC, PLF jettison canbe delayed to reduce the heat flux with minor performance loss. After PLF jettison, the RD-180 is throttledup."So it can be a constraint that really originates from the payload and not necessarily from the launch vehicle.

The assumption is wrong and therefore that isn't the second source.It only affects fairing jettison time. The trajectory isn't really affected.

Jettison time controls when the throttle down to 2.5 G occurs, which is a pretty significant event during booster flight. I guess you can split hairs about what the conversation is about and what is meant by "constraint", but I thought it would be helpful to point out a case where significant events in a trajectory are based off of heating. After all, it seemed to me the question was looking for more sources as to why "heat flux is an important constraint when simulating launcher trajectories".

Where do you see it as a constraint? As free molecular heating?For Pegasus, it is not just the fairing. The wings char.Also, there is interference heating.

Well I just had one study that claimed that heat flux is an important constraint when simulating launcher trajectories. However, I cannot find a second source, so perhaps my assumption is wrong?

A second source would be the Atlas V User Guide:"For Atlas V 500 series missions, the PLF is jettisoned during the booster phase of flight. Before PLF jettison,the RD-180 engine is throttled down to maintain 2.5 g acceleration. Typically, the PLF is jettisoned when the3-sigma free molecular heat flux falls below 1,135 W/m2 (360 Btu/ft2-hr). For sensitive SC, PLF jettison canbe delayed to reduce the heat flux with minor performance loss. After PLF jettison, the RD-180 is throttledup."So it can be a constraint that really originates from the payload and not necessarily from the launch vehicle.

The assumption is wrong and therefore that isn't the second source.It only affects fairing jettison time. The trajectory isn't really affected.

Jettison time controls when the throttle down to 2.5 G occurs, which is a pretty significant event during booster flight. I guess you can split hairs about what the conversation is about and what is meant by "constraint", but I thought it would be helpful to point out a case where significant events in a trajectory are based off of heating. After all, it seemed to me the question was looking for more sources as to why "heat flux is an important constraint when simulating launcher trajectories".

That only applies to the 5m version of Atlas. The 4m fairing is jettison during second stage flight and there is no throttling of the second stage engine. Same applies for the Delta IV and II.

2. If so, are there any engineering solutions that work in the vacuum of space that are commonly applied?

1. Electric propulsion that relies on accelerating ions carries away a positive charge, thus charging the spacecraft with a negative charge that must be dissipated.

2. There are "spacecraft neutralizers" (e.g., the LISA probes have them) that consist of field emitter arrays--arrays of tiny nanospikes that use field emission of electrons to emit electrons directly into the vacuum to neutralize the spacecraft.

1. Electric propulsion that relies on accelerating ions carries away a positive charge, thus charging the spacecraft with a negative charge that must be dissipated.

The exhaust beam of electric thrusters is neutralized, otherwise the positive ions would eventually halt, reverse course, and start accelerating back toward the spacecraft.

The solar wind plasma serves as a ground; thus, I would guess that any positive ions would be neutralized by solar wind electrons, leaving the spacecraft with a net negative charge--hence the need for spacecraft neutralizers. However, I'm not an expert, which is the reason I asked the question in the first place.

Logged

"When once you have tasted flight, you will forever walk the earth with your eyes turned skyward, for there you have been, and there you will always long to return."--Leonardo Da Vinci

1. Electric propulsion that relies on accelerating ions carries away a positive charge, thus charging the spacecraft with a negative charge that must be dissipated.

The exhaust beam of electric thrusters is neutralized, otherwise the positive ions would eventually halt, reverse course, and start accelerating back toward the spacecraft.

The solar wind plasma serves as a ground; thus, I would guess that any positive ions would be neutralized by solar wind electrons, leaving the spacecraft with a net negative charge--hence the need for spacecraft neutralizers. However, I'm not an expert, which is the reason I asked the question in the first place.

I should have been more clear. I was making a statement of fact, based on previous work at a company which designs and manufactures gridded ion and Hall effect thrusters. A hot cathode, integral to the thruster, serves as an electron source, neutralizes the ion beam, and completes the virtual circuit. There is no reliance on in-space electron sources, and this is why the thrusters generate thrust within a ground-based vacuum chamber.

The cathodes are like the Hollow Cathodes depicted here. These can be seen in the Hall Thrusters here, at the top of the devices (often dual-redundant in the Russian designs, as shown).

I've been wondering about a full stage CH4/LOX engine. I always thought that on a FRSC version, you could use an expander cycle to get as many starts as you'd want. But that made me think, can you make a FSC with a single shaft turbopump? Can the turbopump have two turbines on the same shaft?So you can start it up on the expander cycle and then transition to the power of both gas generators. Is doing two separated turbopumps more efficient?I understand that in the H2/LOX case it might get different given the widely different volume, density and specific heat, right?

Our planet rotates on itself and revolves around the sun. If we follow the idea that our galaxy is also in motion, are we not in reality moving at tremendous amounts of speed?.......If we were to create a model of our galaxy and present it in a classroom (for the naked eye)...and we were to move the model over even a fraction of a sliver every minute or hour, this movement would represent millions of kilometers per minute or per hour.......